Gas turbine construction

ABSTRACT

A gas turbine consisting of alternate stators and rotors so fabricated that all stator vanes, all rotor blades and all blade tip seals may be removed through the engine case without requiring disassembly or translation of major portions of the gas turbine engine.

This is a division, of application Ser. No. 601,564, filed Aug. 1, 1975(now U.S. Pat. No. 4,011,718).

BACKGROUND OF THE INVENTION

1. Field of Invention

This invention relates to gas turbine engines and more particularly tosuch engines which have turbine assemblies so constructed that, with aportion of the engine case removed, all airfoil members of the turbinemay be removed without disassembling other engine parts.

2. Description of the Prior Art

While there are patents in the prior art which teach gas turbineconstruction in which combustion chambers, and some select portions ofthe turbine can be removed without disassembling the remainder of theengine, it is believed that the prior art does not teach a constructionin which all airfoil members and blade tip seals of the turbine can beremoved without disassembly of any other engine parts.

SUMMARY OF THE INVENTION

It is the object of this invention to provide a turbine for a gasturbine engine which is so fabricated that all turbine airfoil vanes,blades and blade tip seals may be removed through an access hole inengine case without disassembling any other portion of the engine.

In accordance with the present invention, all airfoil parts and seals soremoved can be replaced utilizing a reassembly procedure which is theexact reverse of the disassembly procedure, with disassembly andreassembly permissible from a forward access position. The presentinvention teaches a turbo machinery stator construction which preventsvane twisting during operation and which minimizes parasitic leakageacross the stator. This stator construction also provides minimumeccentricity between the cooperating seal members carried by the statorand its associated rotor. This stator construction is also of amulti-piece construction so that no single part is required to withstandsubstantially different temperatures in different areas thereof and sothat materials suitable for the particular temperature range ofoperation can be chosen for each part for maximum performance withoutunduly stressing the stator parts. The stator construction may befabricated of minimum tolerance to insure minimum parasitic leakage.

In accordance with a further aspect of the present invention, allturbine airfoil parts are fabricated and supported so that they may beremoved from the engine by freeing them and moving them axially forward.

In accordance with still a further aspect of the present invention, mostturbine airfoil parts and their support assemblies can be removed fromthe engine by freeing them from rearward access position and moving themaxially rearward.

Other objects and advantages of the present invention may be seen byreferring to the following description and claims, read in conjunctionwith the accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is a side view of a typical gas turbine engine with free turbinewhich could utilize the turbine of this invention.

FIGS. 2 and 2A are enlarged cooperating showings through the turbinesection of the FIG. 1 engine.

FIG. 3 is a showing through section line 3--3 of FIG. 2.

FIG. 4 is a showing through adjacent inboard platforms of first stagevanes 52 to show the construction of the feather seal.

FIG. 5 is an enlarged showing of a portion of the connection between thesecond stage stator vanes and its diaphragm member in the second stagestator assembly.

FIG. 6 is a view taken along the line 6--6 of FIG. 5.

FIG. 7 is a perspective showing of the inner end of the second stagevanes to show their lug constructions.

DESCRIPTION OF THE PREFERRED EMBODIMENT

Referring to FIG. 1 we see gas turbine engine 10, which is preferably ofthe free turbine type and consists basically of a gas generator portion12 comprising either an axial flow or centrifugal compressor 14, acombustion section 16, a turbine section 18 and a free turbine 20.Engine 10 is enveloped within case 22, which is preferably of circularcross section and concentric about engine axis or centerline 24. Inconventional manner, air enters inlet 26 of case 22, is compressed inpassing through compressor section 14, is heated and has energy addedthereto in passing through burner 16, has sufficient energy extractedtherefrom in passing through turbine section 18 to drive compressor 14,and is then discharged through free turbine 20. Free turbine 20 ismechanically connected to drive a drive shaft which may either extendfrom the rear end 28 of engine 10 to drive any work generating device,such as an electric generator or a gas pump, or the drive shaft mayextend forwardly and out through the front or forward end 30 of engine10 to drive an engine propeller or be connected to appropriate shaftingto drive a helicopter rotor or other rotatable mechanism, the latterarrangement may be of the type shown in U.S. Pat. No. 3,823,553.

Referring to FIGS. 2 and 2A we see turbine section 18 in greaterparticularity. Basically turbine section 18 includes high pressureturbine rotor assembly 32 and low pressure turbine rotor assembly 34,each of which is mounted from shafting such as 36 and 38 for rotationabout engine centerline or axis 24. Turbine rotor assemblies 32 and 34include a disc member 40 and 42 which carries a plurality of turbineblades 44 and 46 about the periphery thereof, respectively. First andsecond stage stator assemblies, 48 and 50, are mounted concentricallyabout centerline 24 so that first stage stator assembly 48 is positionedforward or upstream of the high pressure turbine rotor assembly 32 andthe second stage stator assembly 50 is positioned between rotorassemblies 32 and 34. Stator assembly 48 includes a plurality of firststage vanes 52 positioned circumferentially thereabout, while statorassembly 50 includes a plurality of second stage vanes 54 positionedcircumferentially thereabout. Blades 44 and 46 and vanes 52 and 54 areairfoil members positioned in annular passage 56 in which the heatedgases from the burner section 16 pass through turbine section 18 enrouteto free turbine 20.

Turbine section 18 is enveloped, in part, within cylindrical case member58 and turbine case 110, which forms parts of engine case 22. Case 58has a forward circumferential flange 60, which is bolted to the forwardportion of case 22 by bolt member 62 and a rearward circumferentialflange 64 which is bolted to the forward section of case 110 by boltmember 66. Flanges 60 and 64 project radially outwardly and inwardly,respectively, and cooperate with the remainder of case 22 so that withbolt members 62 and 66 removed, case 58 may be slid axially rearward toprovide ready access between surfaces 68 and 70 to the interior ofburner section 16 and forward access to turbine section 18 through space69.

As used herein, the term "forward access" means access from a stationforward thereof in the engine and "rear access" means access from astation rearward thereof in the engine.

Transition ducting 72 conducts the heated gases from combustion section16 to annular passage 56 of turbine section 18 and is supported at itsforward end in conventional fashion (not shown) to the conventionallyremovable combustion cans or annular chamber of combustion section 16and is supported at its after end by transition section outer supportring 74 and transition section inner support ring 76. Bolt and nutmembers 78 join transition section 72 to support member 74 and arereadily accessible with case 58 removed. Bolt and nut member 80 connectstransition section 72 to support ring 76 and are readily accessible withcase 58 removed through access ports or openings (not shown) positionedcircumferentially about transition section 72. Transition duct 72 may beof the type shown in U.S. Pat. No. 2,702,454; 2,848,191; or 3,086,363.

First stage stator assembly 48 includes frusto-conical support ringmember 82 which supports the inner end or root of vanes 52 in thatflange member 84, which projects radially inwardly from each vane 52abuts circumferential flange 86 of member 82 and has bolt holes inalignment therewith so that bolt and nut members 88 may extendtherethrough to support the inner end of vanes 52. Segmented retainerplate member 90 extends a selected circumferential distance alongflanges 84 and 86 so that a given nut-bolt unit 88 may position aselected number of vanes 52 due to the action of circumferentiallysegmented retainer plate 90. Member 48 is supported from support ring82, which may in turn be supported from any rigid support (not shown),through ring member 94, to which member 48 may be integrally connectedor may preferably be connected through connecting ring 96 by means ofbolts 98 and nut-bolt arrangements 100 so that member 48 may bedisplaced forward to gain access to sideplate 184. Support member 94 is,in turn, supported from support ring 92 by bolt and nut members 102.Bolt-nut members 104 connect support ring 106 from the root or inner endof vane 52 so that transition duct support ring 76 is supportedtherefrom through slip joint 108.

The outer ends or tips of vanes 52 and 54 are generally supported fromturbine case 110, which is, in turn, connected to engine case 22 bybolt-nut members 66 and 112. Turbine inner case 114 is supported fromturbine case 110 by bolt-nut arrangements 116, 133 and 226 and serves tosupport the outer end or tip of vanes 54 in a manner to be describedhereinafter. Bulk head seal ring 118 is supported from turbine case 110by bolt-nut members 120 and supports the forward portion of inner case114 through nut-bolt arrangements 124. Nut-bolt arrangement 126 supportsfirst vane outer support ring 128 and transition duct outer support 74from support ring 122 on case 114. The outer end of each vane 52 isconnected to support ring 128 by bolt-nut members 130 which extendthrough aligned holes in support ring 128 and radially outwardlyextending flange 132 projecting from the tip of vane 52. A segmentedcircumferentially extending retaining plate 134, comparable to retainingplate 50, and preferably extending equal circumferential arcs therewith,is utilized so that each nut-bolt 130 retains the outer end of aselected plurality of multiplicity of vanes 52.

Second stage vanes 54 are supported at their outer ends in cantileverfashion from turbine inner case 114 which is shaped to define axiallyextending, circumferential recess 140 shaped and sized to receiveaxially extending toe member 142 of the outer platform 144 of each vane54. Case 110 also supports support ring 146 through the action ofbolt-nut member 116 and 226 and cooperates with case 114 to defineaxially extending, circumferential recess 148 to receive toe member 150of vane outer platform 144.

Diaphragm ring 152 of assembly 50 is supported from the inner platform154 of vanes 54 due to a spline connection between platforms 154 andmember 152, as best shown in FIG. 3. Inner platform 154 also includesradially inwardly projecting member 160, which are received in matingengagement in circumferential recess 162 in ring member 164, which isconnected to conical member 152 through bayonet connection 168. Thepurpose of ring member 164 is to avoid the stress which would be createdif diaphragm member 152 were made as a single piece in view of the factthat the various portions of a one-piece diaphragm 152 would beoperating in substantially different temperature zones. With thetwo-piece construction, mainly members 152 and 164, these parts can befabricated of material which is capable of withstanding the temperaturein the zone at which it must operate, part 164 being subjected to muchhigher temperatures and without undue stress due to differential thermalmovements. The stator construction in this area will be betterunderstood by viewing FIGS. 5, 6, and 7.

First viewing FIG. 5, we see ring member 164 connected to diaphragmmember 152 through bayonet connection 168. Bayonet connection 168 is ofconventional design in that spaced tooth members 310 project radiallyoutwardly from diaphragm ring 152 and are spaced so that radiallyinwardly projecting spaced tooth members 312 of ring 164 can be movedaxially between teeth 310 until a continuous, circumferential sealingsurface is generated by the abutment of surfaces 314 and 316, and thenring 164 can be rotated about axis 24 to bring teeth 310 and 312 intosubstantial axial alignment and hence lock ring 164 and diaphragm 152with respect to each other axially.

Ring 164 also includes ring 318 projecting axially forward from surface316 and extending concentrically around axis 24. Ring 318 has aplurality of radially extending slots 320 (see FIG. 6) in spacedcircumferential relationship or array thereabout, while ring portion 322of diaphragm 152 has corresponding radially outwardly opening slots 324in the outer periphery thereof so that ring 164 can be rotated to bringslots 320 and 324 into axial coincidence and serve to form acircumferential array of axially extending female cavity in which toreceive an axially extending lug 326 (see FIG. 7) projecting inwardlyfrom the forward end of vane 54. Vane 54 also has circumferentiallyextending lug 328 at its aft end which is received in circumferentiallyextending, radially opening slot 162 in ring 164. It will therefore beseen that with ring member 164 and diaphragm 152 assembled as shown inFIGS. 5 and 6, each vane 54 can be connected thereto by having vane lugs326 and 328 received in slot combination 324-320 and slot 162,respectively. This serves to lock ring 164 from rotation with respect tovanes 54 and diaphragm 152 and therefore locks the bayonet connection168.

There are several advantages to be gained by the stator construction ofFIGS. 5-7 just described. One such advantage is the fact that thecircumferentially extending vane lug 328 is snuggly received incircumferentially extending ring groove 162, which prevents twisting ofthe vane and thereby ensures that it is always in proper aerodynamicorientation in gas passage 56 and also prevents parasitic leakagebetween surfaces 314 and 316 from areas 330 to 332 (see FIG. 2). Inaddition, in view of the fact that the plurality of axially extendinglugs 326 on the forward end of vanes 54 being received in aligned slots324-320, which is actually a circumferential spline connection betweenvanes 54 and members 164 and 152, this circumferential spline connectionwill provide for minimum eccentricity between the rotor portion 170a andthe stator portion 170b of seal 170. This stator construction, whichincludes the two separate pieces 152 and 164 operating in substantiallydifferent temperature zones, will permit fabricating piece 164 from adifferent material than piece 152, thereby permitting a selection ofmaterials which will provide the necessary temperature tolerance and yetprovide structural integrity and lightness of weight. If diaphragmassembly 152 were one piece, i.e., if members 152 and 164 were integral,this advantage would not be possible and substantial stress would ensuredue to relative thermal movement. Further, since dimension E shown inFIG. 6 between mating surfaces 314-316 and 168a and 168b is minimum inthis design, differential thermal growth between parts 164 and 152 in anaxial dimension is of minimal concern from a stress standpoint and, inview of the fact that the aforementioned mating surfaces between parts162 and 164 are radially extending, relative radial thermal motionbetween parts 152 and 164 is permitted in this design without stressingeither part and while maintaining the necessary structural relationshiptherebetween. It will be realized if substantial relative axial thermalmotion occurred between parts 152 and 164, the aforementioned parasiticleakage would occur in that surfaces 314 and 316 would be drawn out oftheir normal complete circumferential, sealing contact. In addition, thesmall axial dimension D of groove 162 (see FIG. 6) and the correspondingdimension of vane lug 328 minimizes the effect of relative axial thermalexpansion between the vane 54 and ring 164 so that the mating of tab 328in slot 162 further serves to prevent the aforementioned parasiticleakage between areas 330 and 332. This stator design provides theadditional advantage that the parts involved are easy to machine in thatthe walls of slot 162 and surfaces 168a and 316 in ring 164 areparallel, as are surfaces 322a, 316, 168b and 310a in diaphragm 152.

Ring member 152 supports the fixed portion of air seal assembly portion170 at its inner end, due to bolt-nut connection 173.

While rotor assemblies 40 and 42 are shown to be supported for separaterotation by shafts 36 and 38, and this is the preferred arrangement todrive a twin spool, axial flow compressor, it should be borne in mindthat these rotors could be supported and driven in any preferred mannersuch as in unison, for counter-rotation, or any preferred arrangementwithout departing from the spirit hereof.

Shaft 36 is connected by spline connection 171 to shaft 172, which leadsdirectly or indirectly to drive compressor section 14. Nut member 174threadably engages shaft 172 at thread connection 176, and clamps shaft36 at surface 179 on nut 178.

Disc 40 of rotor 32 includes a plurality of conventional circumferentialfir-tree or other type blade receiving slots in its outer periphery sothat matingly shaped fir-tree or other shaped roots of blades 44 may beaxially slid into each such fir-tree slot in disc 40 as shown, forexample, in U.S. Pat. No. 2,686,656 or 2,801,704. A conventional bladedamper 188 is positioned between the inner platform 182 of adjacentblades 44 and bears thereagainst for vibration damping purposes. Theroots of blades 44 and dampers 180 are axially positioned in disc 40between front sideplate 184 and rear sideplate 186. Bolt or nut-on-studmembers 188 and 190 retain the sideplates to the disc front and rearsurfaces adjacent the blade roots, while studs 192 extend through thesideplates for retention purposes at a radially outboard station.Sideplates 184 and 186 are preferably circumferentially segmented toretain a selected number of blades 44 and for ease of assembly anddisassembly.

Second stage blades 46 are similarly attached to the outer periphery ofdisc 42 of rotor assembly 34 through a fir-tree or other conventionalconnection between the roots of blades 46 and corresponding rootreceiving slots in the periphery of disc 42 so that blades 46 may beaxially slid into position and then retained in position by segmentedsideplates 194 and 196, which are retained on studs 198 and 200 by nuts202 and 204. Stud 206 performs a similar function to stud 192, and bladevibration dampers 208 are positioned between the root platforms 205 ofadjacent blades 46 and are axially retained between circumferentiallysegmented sideplates 194 and 196.

Rotors 32 and 34 may be of the type shown in U.S. Pat. No. 3,455,537;3,666,376 or U.S. patent application Ser. No. 455,838 filed Mar. 28,1974 on Gas Turbine Constructions by Kozlin et al and sideplates 184,186, 194 and 196 may be of the variety shown in U.S. Pat. Nos.2,494,658; 2,998,959; or 3,300,179.

It is an important feature of this construction that with case 58unbolted and moved axially, all airfoil members including vanes 52 and54 and blades 44 and 46 may be removed from the engine for repair,replacement or other needed maintenance through the forward access spaceprovided by the translation of case 58. The removal of these airfoilmembers in this fashion will now be described.

By removing bolts 62 and 66, case 58 may be translated axially rearwardto provide an access space 69 between surfaces 68 and 70. Inconventional combustion chamber design (not shown) the combustionchambers can be either removed from the engine either from the accesshole 69 or slid forward within the engine case interior chamber to giveaccess to transition duct 72. By the removal of bolt-nut members 78 and80, the segmented transition duct 72 can be removed through the accesshole or translated forward. With transition duct 72 removed ortranslated forward, we now wish to remove one or more of the first stagevanes 52. To do this, bolts 126 are unbolted and the transition ductouter support case 74 is slid forward into the cavity of combustionsection 16. Support ring 76 and seal ring 108 are also displaced forwardinto cavity of combustion section 16. The removal of bolts 104 and 105permit the removal of first vane inner air seal 106 and outer air seal107, respectively. The removal of seals 106 and 107, which may be ringsor segmented rings, permits the axial removal of the inner and outerfeather seals 109 and 111 from their capture slots 117 which are definedby abutting tip and root platforms 113 and 115 of adjacent first stagevane 52 and extend for substantially the full axial dimension thereof. Ashowing of a typical feather seal is shown in FIG. 4. Feather seals 109and 111 prevent gas leakage between adjacent vanes 52. We then removenuts 88, which frees the segmented retainer plate 90 for removal andhence frees a selected number of vanes 52 at their inner ends. Byremoval of nuts 130 and the segmented retaining plate 134 held inposition thereby, we free the outer ends of vanes 52. Vanes 52 may thenbe removed axially forward and through the access hole 69. In thisfashion, vanes 52 can be removed individually as desired, or in anydesired numbers, or in their totality by merely removing nuts 88 and 130and all retaining plates 90 and 134.

In the alternative, rather than removing vanes 52 individually, we couldhave removed nuts 100 and 130 to permit conical support member 48 andall of the vanes 52 attached thereto to be slid axially forward into thecavity of combustion section 16. Whether we remove all vanes 52individually or we slid all vanes 52 axially forward in attachment tosupport member 48, we bare the blades 44 of the high pressure turbine 32for removal. With vanes 52 so removed, the removal of first stageturbine blades 44 will now be described.

To remove the blades 44, it is first necessary to detach as previouslydescribed, support ring 48 and translate it forward into the combustionchamber cavity. To free the roots of blades 44, segmented sideplate 184is removed. To accomplish this, nuts 188 and 189 are first removed topermit the removal of the segmented sideplate 184. Blade vibrationdampers 180, which may be of any conventional design extending betweensideplates 184 and 186 and bearing against the under surface of the rootplatforms 182 of blades 44, can then be slid axially forward frombetween blades 44. With the damper so removed, the individual blades 44may then be slid axially out of their fir-tree or other root attachmentto disc 40. Stud 192 is securely held in disc 40. With blades 44 soremoved, we will now proceed to remove vanes 54.

With vanes 52 and blades 44 so removed, if we are to remove outer airseals 129, feather seals 131 and vanes 54, we must first remove firstvane outer support ring 128, then remove circumferentially segmented airseal 129. This bares for removal outer vane feather seals 131, which aresimilar in construction to the feather seals shown in FIG. 4. Next,bolts 124 and 120 are removed to free the forward bulk-head seal ring118 and permit it to be axially translated forward into the combustionchamber compartment. Nuts 133 are then removed from the nut-bolt unit116 to free seal plate segments 135 to permit the translation of theinner turbine case 114 forwardly into the combustion chambercompartment. Bolts 137 are then removed. Vanes 54 are displaced axiallyforward slightly with the inner diaphragm member 152 still attached, tofree platform 144 from air seal 250 and vanes 54 are then slid radiallyoutwardly within the spline connection of FIG. 3 to permit the removalof the vane inner feather seals 139, which may be of the FIG. 4 variety.The second stage vanes 54 may then be removed individually.

To remove blades 46, we first slide diaphragm member 152 axially forwardafter first removing bolts 141 and segmented air seal ring 143.Segmented sideplate 194 is next removed by removing nuts 202 and 203.The second blade damper 208, which may be the same type damper as 180for blades 44, is then slid forwardly for individual removal. Blades 46may then be slid axially forward from their firtree or otherconventional attachment to disc 42. If, to permit the removal of blades46, additional axial clearance is needed forward thereof, ring member164 may be disconnected from diaphragm 152 by the action of the bayonetjoint 156 therebetween and moved out of axial alignment with blades 46.The second stage blade outer air seal 250, which is circumferentiallysegmented, may be removed by sliding it axially forward to disengagehooks 252 and 254, then removing the individual segments of seal 250through their ship lap connection.

It will therefore be seen that through the access space 69 created bythe removal of case 58, forward access is provided to turbine section 18and all of the airfoil members 52, 44, 54, and 46 of turbine section 18can be removed and replaced without disturbing any other portion ofengine 10.

It will be evident to those skilled in the art that reassembly ofairfoil members 46, 54, 44, and 52 and seals 111, 129, 131, and 250 maybe accomplished by reversing the procedure just described for theirdisassembly.

As stated earlier, many of the parts of turbine 18 can be removed fromengine case 22 when rearward access is available thereto as now to bedescribed.

The first step in rearward disassembly of turbine 18 would be theremoval of nut 178. This permits the removal as a unit of shaft 36 andsecond stage or low pressure turbine rotor assembly 34 due to thepresence of spline connection 171. It will be evident that if we merelywished to remove one or more of blades 46, sideplates 196 could havebeen removed by the removal of nuts 204 and 206 to permit removalrearwardly of one or all of blades 46.

With low pressure turbine rotor 34 removed as described, nuts 226 can beremoved to thereby permit the axially rearward removal of outer air seal250, outer air seal support 146, and second stage stator assembly 50 andfirst stage outer air seal 129 as a unit, for later disassembly to anydesired extent as previously described.

Next, if first stage turbine rotor assembly 32 is to be removed, tablocks 302 can be opened to permit the removal of bolts 300 and therebypermit the axially rearward removal of the first stage turbine rotorassembly 32. If, in the alternative, we merely wish to remove one ormore blades 44 from rotor assembly 32, this can be done by eitherforward or rearward access. The forward access removal was previouslydescribed and the rearward axis removal would entail the removal of nuts190 and 400, sideplate 186 and then after removing dampers 180, theremoval of the individual blades 44.

If first stage vanes 52 and the other portions of first stage statorassembly 42 are to be removed, case 58 must be released and movedaxially rearwardly as previously described and vanes 44 and the otherportions of stator assembly 48 removed through access hole 69 aspreviously described.

Although the invention has been shown and described with respect to apreferred embodiment thereof, it should be understood by those skilledin the art that other various changes and omissions in the form anddetail thereof may be made therein without departing from the spirit andthe scope of the invention.

Having thus described a typical embodiment of my invention, that which Iclaim as new and desire to secure by Letters Patent of the United Statesis:
 1. A turbomachinery stator assembly concentric about an axis andincluding:(1) a plurality of radially extending airfoil shaped vanespositioned in circumferential array about said axis. (2) means tosupport the radial outer end of said vanes, and (3) means to positionthe radial inner end of said vanes and be supported therefromincluding:(a) a diaphragm member extending radially inwardly from theradial inner ends of said vanes, (b) a spline connecting said diaphragmmember to said vane inner ends so as to conentrically support saiddiaphragm member from said vane inner ends, (c) a circumferentiallyextending ring member positioned radially inward of said vanes, (d)means connecting said ring member to the inner end of said ends of saidvanes so as to axially position said ring member and support said ringmember concentrically about the axis, (e) a bayonet connectionconnecting said ring member to said diaphragm member to lock saidmembers in axial relationship.
 2. A stator assembly according to claim 1wherein said ring member and said diaphragm member are shaped to presentmating surfaces which are in continuous abutting relationshipcircumferentially about said axis to seal against gas flow between saidring member and said diaphragm member.
 3. A stator assembly according toclaim 2 wherein said connecting means between said vane inner ends andsaid ring member constitutes a circumferentially extending grooveopening radially outwardly in said ring member and a matingcircumferentially extending lug member projecting radially inwardly fromthe inner ends of each vane so that all of said circumferentiallyextending lug members are matingly received in said groove to axiallyposition the ring member with respect to the vanes and to concentricallyposition the ring member about the axis.
 4. A stator assembly accordingto claim 3 wherein said spline connection between the inner end of saidvanes and said diaphragm member comprises:(a) a circumferential array ofaxially extending and radially outwardly opening slots in said diaphragmmember, and (b) an axially extending lug projecting from the inner endof each vane and matingly received in said diaphragm slots to form acircumferential spline connection between said vanes and said diaphragmmember to concentrically position said diaphragm member about said axis.5. A stator assembly according to claim 4 and including means to locksaid ring member to prevent rotation thereof about said axis whenassembled.
 6. A stator assembly according to claim 5 wherein said ringlocking means includes a circumferential array of radially opening slotsin said ring member aligning with said slots in said diaphragm memberand matingly receiving said axially extending lugs from said vane innerends in said aligned slots.
 7. A stator assembly according to claim 1wherein said vane outer end support means comprises means to supportsaid vanes in cantilever fashion from said vane outer end support means.8. A stator assembly according to claim 7 wherein said vane outer endsupport means comprises:(a) at least one case member supportedconcentrically about said axis and shaped to define an axiallyrearwardly opening circumferential slot at a forward station and anaxially forwardly opening circumferential slot in an after station, and(b) toe means projecting from said vane outer ends and shaped tomatingly engage said slots to support said vanes therefrom in cantileverfashion.
 9. A stator assembly according to claim 3 wherein the axialdimension of said ring member circumferential slot and said vanecircumferential lug is small to minimize gas leakage between these partsdue to differential thermal motion therebetween.
 10. A stator assemblyaccording to claim 1 wherein said means connecting said vanes to saiddiaphragm member and said ring member permits radial motiontherebetween.
 11. A turbomachinery stator assembly concentric about anan axis and including:(1) a plurality of radially extending vanespositioned in spaced circumferential array concentrically about saidaxis and having:(a) each vane having a circumferentially extending lugand an axially extending lug in axially spaced relation theretoextending radially inwardly from the vane end, (2) means for supportingsaid vanes in cantilever fashion from the vane radially outer ends, (3)a diaphragm member extending radially inwardly from said vane inner endsand including a plurality of radially opening slots spaced incircumferential array thereabout and dimensioned to matingly receivesaid vane axially extending lug so as to concentrically position saiddiaphragm member from said vane inner ends, and (4) a ring memberpositioned radially within the vane inner ends and including a radiallyoutwardly opening circumferential groove shaped to matingly engage thecircumferential lugs at the inner end of each vane so as toconcentrically support and axially position said ring member from saidvane inner ends, and (5) means connecting said ring member to saiddiaphragm so as to axially position said ring member with respect tosaid diaphragm.
 12. A stator assembly according to claim 11 wherein saidconnecting means between said ring member and said diaphragm member is abayonet connection brought into engagement by axially positioning saidring member with respect to said diaphragm member and then rotating saidmembers relative to one another to engage the bayonet connection.
 13. Astator assembly according to claim 12 and including means to preventsaid ring member from rotating with respect to said diaphragm member todisengage the bayonet connection.
 14. A stator assembly according toclaim 13 wherein said ring member lock means includes a circumferentialarray of radially outwardly opening slots in said ring member aligningwith said slots in said diaphragm member to form a series of continuousaxial slots in said ring member and said diaphragm member to jointlyreceive said axially extending lugs from said vane inner ends andthereby prevent rotation about said axis between said diaphragm memberand said ring member.
 15. A stator according to claim 14 wherein saiddiaphragm member and said ring member are shaped to present continuouslymating circumferential surfaces concentric about said axis when soassembled so as to prevent gas leakage therebetween.
 16. A statoraccording to claim 15 and including sealing means supported from theradial inner end of said diaphragm member.
 17. A stator assemblyaccording to claim 11 wherein said vane outer end support meanscomprises:(a) at least one case member supported concentrically aboutsaid axis and shaped to define an axially rearwardly openingcircumferential slot at a forward station and an axially forward openingcircumferential slot in an after station, and (b) toe means projectingfrom said vane outer ends and shaped to matingly engage said slots tosupport said vanes therefrom in cantilever fashion.